Most conventional reinforced composite panel applications are dependent on mechanically fastening at least part of the stiffening elements to the underlying skin. These elements, which in aircraft fuselage applications are known as longitudinal stringers and circumferential frames, are cured in separate processes and later bolted/fastened to a skin shell and in some cases, at least partially to each other. Such concepts, often referred as “black aluminum” due to the similarities with fastened metallic parts, suffer from the low strength resistance of the composites to stress cracks, such as those that may be caused by the necessary drilled holes during fabrication. This characteristic often results in significant weight penalties and the increased costs of assembly operations. Also, the usually high part count associated to these non-integrated concepts increases costs for inventory control, storage, inspection, transportation, and servicing. Although the ideal structural panel may be envisioned to have a part count of one, such integration has been historically avoided due to designs that become overly complex, heavy, or difficult to manufacture.
Many proposals have been made in the art to integrate panels by co-curing, bonding or co-bonding stringers to skins but fewer examples exist on the total integration of these skins to intersecting reinforcements, such as stringers and frames. Such integration usually requires complex tooling designs to allow its disassembly or secondary bonding processes, e.g., bondings that occur after the parts are cured. Such post-cure bonding processes are difficult to certify in primary aircraft composite parts since reliable quality inspection techniques are still lacking, proof testing of each produced part (usually economically unviable) is required and/or over-conservative debonding scenario assumptions must be factored into the component design (thereby also translating into component weight penalties due to the requirement of more robust and/or redundant structural components).
Another challenge to providing a unified one-piece structural panel with intersecting components (e.g., intersecting stringers and frames) relates to the panel geometry. In this regard, most prior attempts to create panels that integrate intersecting reinforcing members can be classified as “grid-panels” as evidenced by U.S. Pat. Nos. 6,110,567, 8,042,315 and 8,079,5491. According to these prior proposals, the intersecting members are laminated at the same level (height) above the skin thereby creating laminating conflicts at their intersections. There also are several disadvantages associated with these prior proposals, including the need for ply cuts, fabric shear while draping the plies, interrupted load paths, labor intensive lamination, and the like. 1 The entire contents of each of these publications, as well as the entire contents of any other publications cited below, are expressly incorporated hereinto by reference.
Closed section stiffeners are especially desirable due to their higher torsional stiffness when compared to open sections. Hat (or omega) stringer concepts, although forming a closed section with the underlying skin in a pristine condition, when delaminated or debonded will behave like open sections—that is, they are unable to stabilize the nearby skins and are unstable themselves since they lack base shell support. In this regard, one proposal (e.g., U.S. Pat. No. 3,995,081) is especially concerned with the potential delaminations of hat stiffeners and adds inner plies to create two “T” flanges at the stiffener base. Other proposals (e.g., U.S. Pat. No. 7,527,222) exist which mechanically fasten the frames over hat stiffener base flanges in an attempt to eliminate such a delamination scenario. Another stringer-only proposal that shares a similar problem (e.g., EP 1800842B1) adopts an inward flanged variation of the hat stiffener.
The use of tubular reinforcements instead of flanged stiffeners (such as hat, omega or inwardly flanged cross-sections) is also known in the art, e.g., from U.S. Pat. No. 4,223,053. According to this prior proposal, a truss-core panel is provided which includes the use of preformed and cured tubes and two separated cure cycles (the first dedicated to produce the tube and the second to crate the entire integrated panel structure. Face sheets may be bonded to the tubes at the same time the face sheets are cured, with a face sheet being overlapped with a tube.
The '053 patent uses the features noted above on a truss-core panel concept. Such a part, however, is difficult to inspect since it creates inaccessible areas; is dependent on the individual manufacturing of a large number of tubes; is difficult to provide interfaces to attach other structures; and significantly reduces the internal structural volume by overlapping tubes. Also, a truss-core panel is structurally not equivalent to a stiffener. Normally, in those composite panels using cores, the cores are used to increase the bending stiffness of the panel by displacing its face sheets, while in composite panels using reinforcements (e.g., stringers and frames), the reinforcements are usually themselves responsible for adding the same bending stiffness.
It would therefore be especially desirable if improvements were provided for integrated composite structural panels which comprise mutually intersecting elongate stiffeners (e.g., stringers and frames). It is towards providing such improvements that the embodiments of the invention disclosed herein are directed.